With reference to FIG. 1, a ducted fan gas turbine engine that can incorporate the invention is generally indicated at 10 and has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, an intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
During operation, air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
Thus it can be said that turbofan gas turbine engines for powering aircraft generally comprise inter alia a core engine, which drives the fan 12. The fan 12 comprises a number of radially extending fan blades 12′ mounted on a fan rotor which is enclosed by a generally cylindrical fan casing.
To satisfy regulatory requirements, such engines are required to demonstrate that if part or all of a fan blade 12′ were to become detached from the remainder of the fan during use (i.e. during rapid rotation of the fan), that the detached part(s) are suitably captured within the engine containment system. Detachment of part of all of a fan blade 12′ in such circumstances is known in the art as a fan blade-off event.
It is known to provide the fan casing with a fan track liner which together incorporate a containment system, designed to contain any released blades parts thereof or associated debris.
FIG. 2 shows a partial cross-section of such a casing and fan track liner, representative of the region indicated by reference numeral 1 in FIG. 1.
In the event of a “fan blade-off” (FBO) event, the detached fan blade 12′ travels radially outward from the fan, and forwards of the fan (i.e. upstream relative to the gas entering the fan). For brevity, hereafter we will refer to a “blade” in the context of a fan blade-off event, but it is to be noted that reference to a “blade” includes not only a detached entire blade, but also a part or fragment thereof, and may also refer to any other significant debris that might be generated by the blade-off event.
The detached blade impacts the fan track liner shown in e.g. FIG. 2 which includes an attrition liner 30, septum 32 and honeycomb layer 34. For example, the blade penetrates the attrition liner 30. It may also penetrate the septum 32 and aluminium honeycomb layer 34 before engaging the hook 36. The fan track liner must therefore be relatively weak in order that any released blade can pass through it essentially unimpeded and subsequently be trapped by the fan casing.
A particular prior art turbomachine casing assembly is shown in FIG. 3. The features described above for FIG. 2 are labelled with the same reference numerals for brevity. As can be seen, in this arrangement, the honeycomb layer 34, for example formed of aluminium, is provided in a tray 38 to be sandwiched between the tray 38 and the septum 32.
Importantly, it can be seen that the septum (and in the particular configuration shown in FIG. 3, the tray 38) are bolted to a lip 40 provided by the hook 36. The bolt and nut assembly 42 is arranged to pinch together the septum 32 (and possibly tray 38) and lip 40. The bolt 43 in the assembly 42 is under tension.
In a fan blade-off event, although the blade will be successfully caught by the arrangement shown in FIG. 3, the action of the detached blade on the fan track liner urges the fan track liner to displace radially outwards. Accordingly, the hook 36 and/or the lip 40 is typically seriously damaged by transferal of the energy of the detached blade from the fan track liner to the hook 36 and/or lip 40 via the bolt 43 due to the presence of holes in the lip.
Such damage may include for example cracking or breaking of the hook 36 and/or lip 40. It may also include for example deformation of the hook 36 or lip 40. Loss of casing integrity in this way is unacceptable in many turbomachines, in particular in gas turbine engines used on aircraft for example. Thus, damage of this sort necessitates repair of the turbomachine, for example by replacing the casing housing the fan blade, and providing the hook 36. This is because of concerns that damage to the hook 36 and/or lip 40 may propagate through the fence and into the casing barrel itself which undermines the integrity of the whole assembly.
Accordingly the present disclosure aims to provide a turbomachine casing assembly, e.g. for a gas turbine engine, that will substantially overcome the problems associated with the prior art assemblies.